Power, Propulsion and

Thermal Design Project

Jesse Cummings

Shimon Gewirtz

Siddharth Parachuru

Dennis Sanchez

Alexander Slafkosky

Mission Itinerary

• Days 1-3: Voyage to moon

• Days 4-7: On the lunar surface

o + 3 Contingency Days

• Days 8-10: Voyage back to Earth

Requirements

• Gross mass 4795 kg

• Must support all mission phases: LEO

checkout, Cis-lunar space, LLO loiter, Lunar

descent and ascent, Lunar surface operations,

Earth EDL.

• Must be capable of limited 6 DOF control

• Must maintain cabin temperatures in Full sun,

Eclipse, Lunar surface dawn/dusk/polar, Lunar

surface 45° sun angle, Lunar surface noon

equatorial

Crew Systems Capsule Design

Selection

• We approached the capsule selection with a

focus on minimizing power requirements and

gross mass.

• With this approach we chose the design

which was 1156.7 kg and which used 116

Watts per day.

• These power and mass requirements are

69% and 9% lower than our second lowest

values.

Crew Metabolic Heat

• We assumed that each crew-day there is a

total of 348 Watts of metabolic heat radiated

per day based on the values from the

ECLIPSE slides in the thermal lecture,

assuming three crew members.

Power:

Requirements

1. Account for two phases of mission

a. In sun-light

b. In darkness

2. Possible Scenarios

a. 13 days of darkness

b. 13 days of light

c. Nominal Case

i. 4 Days Dark on Moon

ii. 7 days Dark on Moon

3. Examine available technologies

4. Select combination to ensure power to system through

energy storage or power generation

Power:

Assumptions and Reasoning

Assumptions

1. Insolation Constant=1394 W/m^2 is actually

constant over the range of the mission

2. Technologies examined can perform to capabilities

3. 100 W of additional power assumed for thruster

solenoids as a conservative estimate.

Reasoning

1. Examine worst case to determine best energy

storage device

2. Examine best case to determine best power

generation device

3. Create a combination to meet requirements

Power:

Selections I

Battery and Cells

1. Ni-Cd

2. IPV

3. CPV

4. Ni-MH

5. Li-Ion

6. NaS

Solar Panels

1. GaAs

2. 2 Junction

3. 3 Junction

4. 4 Junction

5. Single Crystal Si

6. CIGS

Power:

Selections I

These Technologies were excluded:

Type - Reason

• Nuclear and Solar Thermal - Size and Mass

• Flywheel - Low Watt Hours per kg @ SOA

• H2-O2 Fuel Cell - Storage of H2 and O2

• Chemical Thermal - Scalability

Power:

Trade Study on Mass of Energy Cells

Power:

Trade Study on Mass of Batteries

Power: Trade Study on Volume of Battery and Cells

Power:

Comparison of Solar Panels

Power:

Conclusions

Li-Ion Cells:

• Worst case mass = 1298 kg

• Nominal mass (4 Days) = 399.4 kg

• Nominal Volume (4 Days)= 0.25 m^3

Solar Panels:

• Mass = 4.16 kg

• Area= 0.853 m^2

Thermal

• In AU units, difference between Earth and

Moon is significantly small o Moon AU ≈ Earth AU = 1 AU

• Solar Flux Is = 1394 W/m2 (at 1 AU)

• Stefan-Boltzmann Constant σ = 5.67 * 10-8

W/m2K4

• Total Power required for lunar crew module

Pint = 764 W

• Desired cabin temperature T = 293 K (Room

Temperature)

Thermal Calculations

• Used full Stefan - Boltzmann equation and

solved for T (Cabin Temperature)

o As = surface area exposed to sun

o Arad = total surface area

o Tenv = environment temperature

Thermal:

Trans Lunar

• For the trans lunar case, we are assuming

that the solar flux hits everywhere on the

spacecraft expect for the base.

o We make this assumption because this is the max

surface area that the solar flux can hit on the

spacecraft in free space assuming any orientation.

• We also assumed that the environment

temperature in free space is approximately

0° C. o A reasonable assumption to make because

(Cabin temperature)4 >> (Free space temperature)4

Thermal:

Eclipse General

• For the eclipse conditions, we are assuming

that the spacecraft is completely eclipsed

during both the earth and moon orbits.

• During the earth orbit, we assume that the

environment temperature is 280 K based on

the solar flux at the earth's distance from the

Sun.

• During the moon orbit, we assume that the

environment temperature is 0 K because it is

considered to be free space.

Thermal:

Lunar Surface

The Surface of the Moon has different

temperatures at different location and time.

Below is the most extreme temperatures to

design for worst cases.

Lunar Surface Temperature (K)

Dawn 120

Polar 230

Dusk 290

45° Sun Angle 370

Noon Equatorial 390

Thermal:

Lunar Surface

• Lunar crew module will have different

exposure to Sun at different times o Excluding the bottom surface area (no sun

exposure)

• At Polar/Dusk/Dawn o 1/3 of surface area exposed

• At 45° Sun Angle

o 1/2 of surface area exposed

• At Noon Equatorial o All of surface area exposed

o Sun is directly above lunar crew module

Thermal:

Coating Material Properties

• For initial thermal calculations, calculated all

cabin temperatures with different coating o Different lunar surface temperatures

o Different emissivities (ε) and absorptivities (α)

Coating Absorptivity (α) Emissivity (ε)

White 0.2 0.8

Black 0.9 0.8

Aluminum 0.3 0.3

Polished Metals 0.2 0.01 ≈ 0

Thermal:

Lunar Surface Trade Study

Thermal:

Emittance

• From our trade study of different coating materials , we found

that aluminum coating (ε = 0.3 and α = 0.3) has the least

variation of cabin temperature from the desired cabin

temperature of 293 Kelvin.

• After doing thermal calculations for all the different conditions

(trans lunar, eclipses, and moon surfaces at different times)

with aluminum coating, we decided to have the lowest

temperature on the trip be equal to the required cabin

temperature of 293 Kelvin, so as to only utilize radiators and

not any heaters.

• For us to meet this condition we found that we would need to

have an effective emittance of 0.056

Insulation Trade Study

Thermal:

Multi Layered Insulation

• We were able to achieve this condition by adding 5

layers of an 850-3M Mylar-Aluminum Backing

insulation with aluminum coating on both sides to get an

effective emissivity of 0.0557 so as to get the cabin

temperature during the coldest situation, during the

moon eclipse, to be 293 Kelvin.

• Similar to the effective emissivity, the effective

absorptivity of aluminum decreases with more layers.

• Final absorptivity α = 0.05 and emissivity ε = 0.0557

Thermal:

Cabin Temperatures

• Cabin temperatures in different situations

with multi layered insulation (MLI)

Situation Temperature (Kelvin) Temperature (° Celsius)

Lunar Surface: At Dawn 387 114

Lunar Surface: At Dusk 353 80

Lunar Surface: At Polar 367 94

Lunar Surface: At 45°

latitude

430 157

Lunar Surface: At equatorial 464 193

Trans Lunar (Free Space) 390 117

Eclipse in Earth Orbit 341 68

Eclipse in Lunar Orbit 293 20

Thermal:

Radiator

• Will use Traveling Wave Tube Amplifier

(TWTA) radiators with Optical Solar

Reflectors (OSR) covering them

o Total Mass = 31 kg

o Emissivity of OSR: ε = .77

o Absorptivity of OSR: α = .06

• Radiators will be positioned between top and

middle thrusters

Thermal:

Radiator

• Used the Stefan-Boltzmann equation to

solve for the area required for OSR

depending on the power generated by the

spacecraft.

• Area of radiators ≈ 2 m2

o Radiators will deploy away from the lunar surface at

all times so that the environment temperature is

reduced.

o Designed for the worst case condition that the OSR

panels are facing the sun at all times

o Calculated for the OSR panels area to perform at

room temperature for the cabin.

Thermal:

Position of Radiators

Side View of Radiators Top View of Radiators

(when deployed)

Propulsion:

Requirements

Limited 6 DOF

1. Translational delta V = 50 m/s

2. Attitude Hold in Dead Band for Return

3. Overcome 500 Nm of Aerodynamic

moments due to reentry (Pitch and Yaw)

4. Rotate Spacecraft 180 degrees in 30

seconds (Roll)

Propulsion:

Attitude Control System

The attitude control system will consist of

25 attitude thruster nozzles.

All the nozzles will be recessed into the craft

walls so that they will be protected from

forces on the nozzle walls from drag forces

and heating on re-entry.

The nozzles near the heat shield need to be

placed slightly higher up from the bottom

because of the extreme heating of the heat

shield region.

Attitude Control System Diagram

• 4 thrusters radially at the

top for pitch/yaw

adjustments in conjunction

with radial thrusters on

bottom

• 1 axial thruster at the top

for translational movement

along the z-axis

Attitude Control System Diagram

• 4 thrusters in the x-y plane

around the center of gravity.

For translational motion only

- no moment generated

because of placement.

Attitude Control System Diagram

• 16 thrusters (4 separate groupings)

spaced equally radially. Each

grouping has one radially, one

axially, and two (opposite)

azimuthally.

• Radial thrusters used for pitch/yaw

motion and balance in the moment

they produce, with thrusters on the

top.

• Azimuthal thrusters used for roll

movement.

• Axial thrusters for translational

movement axially with the one on

the top (calibrated to balance the

power of the thruster axially on top.)

Propulsion:

Translational Assumptions

1. delta V = 50 a. The total delta V required for all translational

adjustments

2. Rocket Equation is a sufficient model

3. Mass and volume are primary considerations

4. Power and storage requirements are

secondary considerations

Propulsion:

Translational Selection I

Cold Gas

H2

He

Methane

N2

Air

Argon

Krpton

Freon 14

SOA

NTO-MMH1

1=NASA-Document, See Sources Slide

Propulsion:

Translational Selection II

These Technologies were excluded:

Type - Reason

• Electrical - High Power, Low Thrust

• Nuclear - High Mass

• Solid (Chemical) - Configuration

• Air Breathing (Chemical) - In Space

• Sails - Mission Design

• ED Tether - Mass and Volume

Propulsion:

Translational Reasoning

1. Examine simplest case first (Cold Gas)

1. Examine State of the Art

1. Plot Mass versus Volume for a given system

Propulsion:

Translational Mass versus Volume

Propulsion:

Translational Conclusions I

The best option is the SOA NTO-MMH

This is due to:

• Isp=324 seconds

• Constituent densities at stored temperatures

• Oxidizer (NTO, Nitrogen Tetroxide)

• Fuel (MMH, Monomethyl Hydrazine)

• Earth Storable (liquid at ~290 K)

• Oxidizer to Fuel Ratio=1.65

• Hypergolic (combusts on contact with each other)

• Utilized on the Space Shuttle RCS system

Propulsion:

Translational Conclusions II

For our translational propulsion requirements

we will need:

NTO:

• Mass = 37.74 kg

• Volume = 0.0262 m^3

• Tank (PMD) Mass = 0.052 kg

MMH:

• Mass = 62.26 kg

• Volume = 0.0708 m^3

• Tank (PMD) Mass = 0.086 kg

Propulsion:

Translational Conclusions III

From NASA2 the pressurization system with He

to keep the fuel and oxidizer vapor stable is

about 2 kg He per 55 kg (fuel+oxidizer)

Thus:

• Mass He= 3.64 kg

• @ 200 ATM (Mass Opt. Tank) and 293 K

• Volume He= 1.09 x 10-4 m^3

• Tank Mass He= 3.013 kg

Propulsion:

Dead Band Assumptions

• The amount of acceptable drift in roll, pitch,

and yaw assumed for the dead band drift

during the trip to Earth is 5 degrees in either

direction from the desired path.

• It was also assumed that the lowest burn

time for the thrusters is 0.1 seconds as

limited by the solenoid valves that control the

flow of propellant through the engine.

Propulsion:

Dead Band Reasoning

• An angular velocity can be calculated from

the torque that each set of thrusters delivers

when fired, the moment of inertia about the

axis of rotation, and the lowest burn time

assumed.

Propulsion:

Dead Band Reasoning

• Once an angular velocity can be determined

as the result of an impulse bit delivered from

a set of thrusters, the approximate amount of

fuel consumed to correct the dead band drift

over the entire trip to Earth can be

calculated.

• After the first initial drift, the time before the

next adjustment is completely dependent on

the angular velocity generated by the

impulse bit, which is extremely small.

Propulsion:

Dead Band Data

• The amount of fuel required to correct for

drift during the 3 day trip to Earth is

negligible. The masses of fuel consumed

were less than a gram of propellant total.

• Because the force requirements of the

thrusters are so low as the result of strategic

placement, the mass flow-rate of fuel

through the engines is very small: 34.8 g/s

for the roll thrusters and 44.1 g/s for the pitch

and yaw thrusters.

Propulsion:

Dead Band Conclusion • The conclusion that can be drawn regarding fuel

consumption for corrections to the dead band drift

is that the craft barely drifts at all, and thus

negligible amounts of propellant are consumed.

• This finding can be attributed to the use of a Isp

propellant and a high moment of inertia on the craft,

as well as a very strategic arrangement of thrusters

around the spacecraft.

• Extraneous factors that were not included in this

analysis justify applying a safety factor to the

amount of propellant allocated for the translational

propulsion to ensure enough fuel in case

adjustments must be made to attitude.

Propulsion:

Pitch and Yaw Assumptions

• In order to be able to counteract a moment

of 500 N-m, two coordinated thrusters,

positioned at the top of the craft and along

the bottom, are fired.

• The two thrusters also generate equal

forces, negating any translational velocity

that might be imparted to the craft.

Propulsion:

Pitch and Yaw Reasoning

• In order to calculate the required force to

negate a 500 N-m moment during re-entry,

the two thrusters must deliver combined

moments that equal 500 N-m.

• The forces required to generate those

moments are significantly lower due to the

large moment arm for the thrusters

positioned at the top of the vehicle.

Propulsion:

Pitch and Yaw Data and Conclusion

• When taking this into consideration, each

pitch/yaw thruster needs to be able to deliver

70 N of force.

• This also means that each thruster has a

mass flow-rate of 44.1 g/s, as listed before.

Such a low mass flow-rate is especially

useful for a re-entry situation, where a long,

continuous burn may be necessary to keep

the vehicle stable.

Propulsion:

Roll Assumptions

• The roll thrusters are positioned azimuthally

in a ring as close to the base of the vehicle

as possible without risking damage from re-

entry heating to give the highest possible

moment arm.

• Each roll thruster is also assumed to deliver

the same amount of force.

Propulsion:

Roll Reasoning

• An equation from the lecture notes on

propulsion was used to find the required

torque to achieve the requirement of rotating

180 degrees in 30 seconds or less.

Propulsion:

Roll Reasoning

• By applying a constant torque to accelerate

the roll of the craft to 12 degrees/sec and an

equal and opposite torque to bring the craft

to rest, the vehicle will experience the

smoothest possible acceleration from a

constant burn.

Propulsion:

Roll Data and Conclusion

• The total constant applied torque needed to

execute a 180 degree roll maneuver in 30

seconds is 184.4 N-m. When divided by the

radius of the ring of roll thrusters, the required

force for each thruster is very low: 27.7 N.

• This means that the mass flow-rate for the roll

thrusters is 34.8 g/s of propellant. This is lower

than the mass flow-rate for pitch and yaw

control, and the 180 degree roll maneuver,

when performed over 30 seconds, consumes

half a kilogram of propellant.

Summation Mass Tabulation

System Component Mass (kg) Mass (kg)

Crew Module 1156.7

Power 1302.16

Li-Ion Cell 1298

Solar Panels 4.16

Thermal 30

Radiator 30

Propulsion 170.8

Fuel (MMH) 62.26

Fuel Tank 0.086

Oxidizer (NTO) 37.74

Oxidizer Tank 0.052

RC Nozzles 64

He Mass (at 200 atm) 3.64

He Tank 3.013

Total 2660

Sources

Hutton, R. E. Lunar Surface Models. Tech. no. SP-8023. Washington, D.C: National

Aeronautics and Space Administration, 1969. Print.

Zhongmin, Deng. "Optimization of a Space Based Radiator." Applied Thermal

Engineering31 (2011): 2312-320. Web

Vasavada, Ashwin R. Near-Surface Temperatures on Mercury and the Moon and the

Stability of Polar Ice Deposits. Publication no. Icarus 141. Los Angeles, CA:

Academic, 1999. Print.

"Thermal Control System Design." N.p., n.d. Web. 8 Nov. 2012.

<http://ams.cern.ch/AMS/Thermal/OHB-CERN-100501.PDF>.